On separating the effect of corrosion on inter-lamellar fatigue of thin sheet AA7079-T6

Abstract

AA7079-T6 was widely used in the 1960s for manufacturing aircraft due to its high mechanical strength. Subsequently, this alloy was discontinued due to very poor fracture toughness, crack growth resistance, and corrosion resistance. A large strategic cargo carrier has aft upper fuselage (crown) skin made of this alloy. Various cracks have been identified in this area which has led to significant inspection of the aircraft. Visual inspection of the complete crown skin is done at least three times a year to detect cracks. A detailed failure analysis of several cracks from a few replaced crown skins was conducted to understand the root cause of these cracks. The failure analysis revealed that initial propagation was along the thickness of the skin which then diverged and propagated along mid thickness parallel to the surface of the skin. This crack propagation was accompanied by branching along the grain boundaries in the propagation path. Such crack morphologies are typical of stress corrosion cracking (SCC). However, study of the fracture surfaces revealed the presence of fatigue striations. Comprehensive analysis of the loading condition and operating environment does confirm that fatigue and corrosion occurs separately for these cracks. The fatigue crack propagation occurs while the aircraft is in flight and corrosion occurs while the aircraft is on ground. This corrosion weakens the material preferentially along the grain boundary leading to intergranular fatigue crack growth in flight. For rolled thin sheet product, this intergranular crack appears as an inter-lamellar crack as the grains are oriented in one line. This hypothesis was also verified with controlled corrosionfatigue experiments in the laboratory which replicated the failure surface obtained in service. Experiments were also conducted to determine the threshold for crack growth in fatigue as well as SCC. The fatigue crack growth threshold was observed at 2.2 MPa√m and the SCC threshold was found to be 8.5 MPa√m. The crown skin area where the cracks are located is loaded only during flight when the environment is benign for SCC to occur. On the ground where the environment is amenable to corrosion, there is very little load on the crown skin area where cracks are located. This results in very low stress intensity values and no crack propagation occurs by SCC. The cracks are therefore believed to have propagated during flight by fatigue loading, which is also supported by failure analysis results. © US Government 2009.